The present invention relates in general to rocket engines, and, in particular to a new and useful take-off thrust nozzle for reaction engines, particularly rocket ramjet engines, which have an ablating lining.
In ramjet engines for propelling missiles, so called booster nozzles are known, for example, from German AS No. 1,181,496, which are intended to temporarily reduce the air velocities in the combustion chamber during ignition of the main engine, so as to ensure a satisfactory ignition. Upon starting the ignition process in the combustion chamber, the booster nozzle is ejected as a whole or in parts or fragments. To this end the booster nozzle is made of a material having a sufficiently low fusion point, and the structure of the nozzle is sectioned into individually larger areas by perforation rows. Upon ignition in the combustion chamber, the heat first fuses the webs between the individual perforations so that the booster nozzle is disintegrated into individual fragmentary parts which are expelled rearwardly.
It is further known in rockets and ramjet engines for propelling missiles, to fill the common combustion chamber, which is used for both starting and sustaining the flight, with a solid start fuel for the starting period and to design the following discharge nozzle as a combined sustainer nozzle having a throat structure which is formed by an inside lining of graphite, glass-reinforced plastic, or the like, and serves as the start nozzle. At the end of the starting period, this start nozzle is separated and expelled rearwardly as a whole or in broken parts. Where such missiles are launched from a carrier or launching aircraft the fragments thus produced are very hazardous for the carrier aircraft or other following aircraft.